The present invention relates generally to gas turbine engines, and, more specifically, to turbine shrouds therein.
In a gas turbine engine, air is pressurized in a compressor, mixed with fuel in a combustor, and ignited for generating hot combustion gases which flow downstream through one or more turbine stages which extract energy therefrom. A high pressure turbine (HPT) first receives the combustion gases from the combustor and extracts energy therefrom for powering the compressor. A low pressure turbine (LPT) follows in turn the HPT for extracting additional energy for providing output energy typically used for powering a fan disposed upstream of the compressor in a typical aircraft gas turbine engine application.
The HPT includes a stationary turbine nozzle having a plurality of circumferentially spaced apart stator vanes which control discharge of combustion gases from the combustor. The HPT also includes at least one rotor stage having a plurality of circumferentially spaced apart turbine rotor blades extending radially outwardly from a supporting rotor disk. The blades include airfoils which receive combustion gases from the nozzle and extract energy therefrom for rotating the rotor disk and in turn rotating the compressor. The airfoils are typically hollow and include internal cooling circuits therein through which a portion of pressurized air bled from the compressor is channeled for cooling the blades.
Surrounding the rotor blades is an annular turbine shroud fixedly joined to the surrounding stator casing. The shroud is suspended closely atop the blade tips for providing a small gap or tip clearance therebetween. The tip clearance should be as small as possible to provide an effective fluid seal thereat during operation for minimizing the amount of combustion gas leakage therethrough for maximizing efficiency of operation of the engine. However, due to differential thermal expansion and contraction of the rotor blades and surrounding turbine shroud, the blade tips occasionally rub against the inner surface of the shroud causing abrasion thereof.
Since the blade tips are at the radially outermost end of the rotor blade and are directly exposed to the hot combustion gases, they are difficult to cool and the life of the blade is thereby limited. Furthermore, during a blade tip rub with the surrounding shroud, the blade tips are additionally heated by friction which additionally affects the blade useful life. The friction heat generated during a blade tip rub further increases the radial expansion thereof and correspondingly increases the severity of the tip rub.
Since the turbine shroud itself is exposed to the hot combustion gases, it too is also cooled by bleeding a portion of the pressurized air from the compressor, which is typically channeled in impingement cooling against the radially outer surface of the turbine shroud. Turbine shrouds typically also include film cooling holes extending radially therethrough with outlets on the radially inner surface of the shroud from which is discharged the cooling air in a film for cooling the inner surface of the shroud.
Since blade tip rubs are unavoidable for maximizing performance of the turbine, both the turbine shrouds and blade tips are subject to abrasion wear. However, such abrasion may cause the film cooling holes in the turbine shrouds to plug which can additionally adversely affect the useful life of the turbine shroud.
Accordingly, it is desired to provide an improved turbine shroud for cooperating with turbine rotor blade tips during tip rubs for reducing the severity of the tip rubs and reducing friction heating of the blade tip.